Airfoil fluid curtain to mitigate or prevent flow path leakage

ABSTRACT

Flow path assemblies for gas turbine engines are provided. For example, a flow path assembly comprises an inner wall defining an inner boundary of a flow path and a plurality of pockets therein, and a unitary outer wall defining an outer boundary of the flow path. The unitary outer wall includes combustor and turbine portions that are integrally formed as a single unitary structure. The flow path assembly further comprises a plurality of nozzle airfoils that each have an inner end radially opposite an outer end and define an internal cavity for receipt of a flow of cooling fluid. The inner end of each nozzle airfoil is received in one of the plurality of inner wall pockets and defines an outlet for the flow of cooling fluid to flow from the internal cavity to the pocket, which forms a fluid curtain to discourage fluid leakage from the flow path.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a continuation of and claims priority to U.S.application Ser. No. 15/426,354 (issued as U.S. Pat. No. 10,253,643),filed Feb. 7, 2017, the contents of which are incorporated herein byreference.

FIELD

The present subject matter relates generally to gas turbine engines.More particularly, the present subject matter relates to outer and innerflow path boundary configurations for receipt of stator airfoils anddevelopment of a fluid curtain across the stator airfoils to mitigate orprevent leakage from the flow path.

BACKGROUND

A gas turbine engine generally includes a fan and a core arranged inflow communication with one another. Additionally, the core of the gasturbine engine generally includes, in serial flow order, a compressorsection, a combustion section, a turbine section, and an exhaustsection. In operation, air is provided from the fan to an inlet of thecompressor section where one or more axial compressors progressivelycompress the air until it reaches the combustion section. Fuel is mixedwith the compressed air and burned within the combustion section toprovide combustion gases. The combustion gases are routed from thecombustion section to the turbine section. The flow of combustion gasesthrough the turbine section drives the turbine section and is thenrouted through the exhaust section, e.g., to atmosphere.

More particularly, the combustion section includes a combustor having acombustion chamber defined by a combustor liner. Downstream of thecombustor, the turbine section includes one or more stages, for example,each stage may include a plurality of stationary nozzle airfoils as wellas a plurality of blade airfoils attached to a rotor that is driven bythe flow of combustion gases against the blade airfoils. The turbinesection may have other configurations as well. In any event, a flow pathis defined by an inner boundary and an outer boundary, which both extendfrom the combustor through the stages of the turbine section.

Typically, the inner and outer boundaries defining the flow pathcomprise separate components. For example, an outer liner of thecombustor, a separate outer band of a nozzle portion of a turbine stage,and a separate shroud of a blade portion of the turbine stage usuallydefine at least a portion of the outer boundary of the flow path.However, utilizing separate components to form each of the outerboundary and the inner boundary requires a great number of parts, e.g.,one or more seals may be required at each interface between the separatecomponents to minimize leakage of fluid from the flow path, which canincrease the complexity and weight of the gas turbine engine withouteliminating leakage points between the separate components. Therefore,flow path assemblies may be utilized that have a unitary construction,e.g., a unitary outer boundary structure, where two or more componentsof the outer boundary are integrated into a single piece, and/or aunitary inner boundary structure, where two or more components of theinner boundary are integrated into a single piece.

A unitary construction of the flow path assembly may be furthered byseparating the turbine nozzle airfoils, which also may be referred to asstator vanes, from the outer boundary structure and the inner boundarystructure. As such, the outer boundary structure and/or the innerboundary structure each may be constructed as a unitary structure or maybe constructed together as a single unitary structure, with the nozzleairfoils inserted and secured during subsequent assembly. Separating thenozzle airfoils from the outer and inner boundary structures of the flowpath assembly thereby may simplify manufacturing, as well as reduceinternal stresses compared to flow path assemblies comprising nozzleairfoils that are integral with the outer and/or inner boundarystructures. However, separating the nozzle airfoils from the outer andinner boundaries may introduce points at which, e.g., fluid flowingthrough the flow path can leak from the flow path or where the fluidflowing through the flow path can leak from the pressure side of thenozzle airfoils to the suction side of the nozzle airfoils. This latterform of leakage may be referred to cross-over leakage, and cross-overleakage, from the higher pressure side to the lower pressure side of theairfoils, can negatively impact engine performance.

Accordingly, improved flow path assemblies would be desirable. Forexample, a flow path assembly utilizing a flow of cooling fluid to forma fluid curtain across a nozzle airfoil to discourage cross-over leakagewould be beneficial. Further, providing more than one source of coolingfluid to form the fluid curtain would be advantageous. Additionally, aflow path assembly receiving a flow of cooling fluid that providescooling to a plurality of nozzle airfoils, at least one of an inner oran outer boundary of the flow path, and/or downstream components of theflow path would be useful.

BRIEF DESCRIPTION

Aspects and advantages of the invention will be set forth in part in thefollowing description, or may be obvious from the description, or may belearned through practice of the invention.

In one exemplary embodiment of the present disclosure, a flow pathassembly for a gas turbine engine is provided. The flow path assemblycomprises an inner wall defining an inner boundary of a flow path and aunitary outer wall defining an outer boundary of the flow path. Theinner wall further defines a plurality of pockets therein. The unitaryouter wall includes a combustor portion extending through a combustionsection of the gas turbine engine and a turbine portion extendingthrough at least a first turbine stage of a turbine section of the gasturbine engine. The combustor portion and the turbine portion areintegrally formed as a single unitary structure. The flow path assemblyfurther comprises a plurality of nozzle airfoils. Each nozzle airfoilhas an inner end radially opposite an outer end, and each nozzle airfoildefines an internal cavity for receipt of a flow of cooling fluid. Theinner end of each nozzle airfoil is received in one of the plurality ofpockets defined in the inner wall. Moreover, the inner end of eachnozzle airfoil defines an outlet for the flow of cooling fluid to flowfrom the internal cavity to the pocket, and the cooling fluid flowingfrom the outlet of each nozzle airfoil forms a fluid curtain todiscourage fluid leakage from the flow path.

In another exemplary embodiment of the present disclosure, a flow pathassembly for a gas turbine engine is provided. The flow path assemblycomprises an inner wall defining an inner boundary of a flow path and aunitary outer wall defining an outer boundary of the flow path. Theinner wall further defines a plurality of pockets therein. The unitaryouter wall includes a combustor portion extending through a combustionsection of the gas turbine engine and a turbine portion extendingthrough at least a first turbine stage of a turbine section of the gasturbine engine. The combustor portion and the turbine portion areintegrally formed as a single unitary structure. The flow path assemblyalso comprises a plurality of nozzle airfoils. Each nozzle airfoil hasan inner end radially opposite an outer end, and the inner end of eachnozzle airfoil received in one of the plurality of pockets defined inthe inner wall. A flow of cooling fluid is directed into each pocket,and the flow of cooling fluid forming a fluid curtain to discouragefluid leakage from the flow path into the pocket.

In a further exemplary embodiment of the present disclosure, a flow pathassembly for a gas turbine engine is provided. The flow path assemblycomprises an inner wall defining an inner boundary of a flow path and aunitary outer wall defining an outer boundary of the flow path. Theunitary outer wall includes a combustor portion extending through acombustion section of the gas turbine engine and a turbine portionextending through at least a first turbine stage of a turbine section ofthe gas turbine engine. The combustor portion and the turbine portionbeing integrally formed as a single unitary structure. Further, theunitary outer wall defines a plurality of pockets therein. The flow pathassembly also comprises a plurality of nozzle airfoils, each nozzleairfoil having an inner end radially opposite an outer end. The outerend of each nozzle airfoil received in one of the plurality of pocketsdefined in the unitary outer wall. A flow of cooling fluid is directedinto each pocket, and the flow of cooling fluid forming a fluid curtainto discourage fluid leakage from the flow path into the pocket.

These and other features, aspects and advantages of the presentinvention will become better understood with reference to the followingdescription and appended claims. The accompanying drawings, which areincorporated in and constitute a part of this specification, illustrateembodiments of the invention and, together with the description, serveto explain the principles of the invention.

BRIEF DESCRIPTION OF THE DRAWINGS

A full and enabling disclosure of the present invention, including thebest mode thereof, directed to one of ordinary skill in the art, is setforth in the specification, which makes reference to the appendedfigures, in which:

FIG. 1 provides a schematic cross-section view of an exemplary gasturbine engine according to various embodiments of the present subjectmatter.

FIG. 2 provides a schematic exploded cross-section view of a combustionsection and a high pressure turbine section of the gas turbine engine ofFIG. 1 according to an exemplary embodiment of the present subjectmatter.

FIG. 3 provides a schematic cross-section view of the combustion sectionand high pressure turbine section of FIG. 2 according to an exemplaryembodiment of the present subject matter.

FIG. 4 provides a partial perspective view of a portion of an integralouter boundary structure and inner boundary structure of the combustionsection and high pressure turbine section of FIG. 2 according to anexemplary embodiment of the present subject matter.

FIG. 5 provides a perspective view of a nozzle airfoil of a flow pathassembly of the combustion section and high pressure turbine section ofFIG. 2 according to an exemplary embodiment of the present subjectmatter.

DETAILED DESCRIPTION

Reference will now be made in detail to present embodiments of theinvention, one or more examples of which are illustrated in theaccompanying drawings. The detailed description uses numerical andletter designations to refer to features in the drawings. Like orsimilar designations in the drawings and description have been used torefer to like or similar parts of the invention. As used herein, theterms “first,” “second,” and “third” may be used interchangeably todistinguish one component from another and are not intended to signifylocation or importance of the individual components. The terms“upstream” and “downstream” refer to the relative direction with respectto fluid flow in a fluid pathway. For example, “upstream” refers to thedirection from which the fluid flows and “downstream” refers to thedirection to which the fluid flows.

Referring now to the drawings, wherein identical numerals indicate thesame elements throughout the figures, FIG. 1 is a schematiccross-sectional view of a gas turbine engine in accordance with anexemplary embodiment of the present disclosure. More particularly, forthe embodiment of FIG. 1, the gas turbine engine is a high-bypassturbofan jet engine 10, referred to herein as “turbofan engine 10.” Asshown in FIG. 1, the turbofan engine 10 defines an axial direction A(extending parallel to a longitudinal centerline 12 provided forreference) and a radial direction R. In general, the turbofan 10includes a fan section 14 and a core turbine engine 16 disposeddownstream from the fan section 14.

The exemplary core turbine engine 16 depicted generally includes asubstantially tubular outer casing 18 that defines an annular inlet 20.The outer casing 18 encases, in serial flow relationship, a compressorsection including a booster or low pressure (LP) compressor 22 and ahigh pressure (HP) compressor 24; a combustion section 26; a turbinesection including a high pressure (HP) turbine 28 and a low pressure(LP) turbine 30; and a jet exhaust nozzle section 32. A high pressure(HP) shaft or spool 34 drivingly connects the HP turbine 28 to the HPcompressor 24. A low pressure (LP) shaft or spool 36 drivingly connectsthe LP turbine 30 to the LP compressor 22. In other embodiments ofturbofan engine 10, additional spools may be provided such that engine10 may be described as a multi-spool engine.

For the depicted embodiment, fan section 14 includes a fan 38 having aplurality of fan blades 40 coupled to a disk 42 in a spaced apartmanner. As depicted, fan blades 40 extend outward from disk 42 generallyalong the radial direction R. The fan blades 40 and disk 42 are togetherrotatable about the longitudinal axis 12 by LP shaft 36. In someembodiments, a power gear box having a plurality of gears may beincluded for stepping down the rotational speed of the LP shaft 36 to amore efficient rotational fan speed.

Referring still to the exemplary embodiment of FIG. 1, disk 42 iscovered by rotatable front nacelle 48 aerodynamically contoured topromote an airflow through the plurality of fan blades 40. Additionally,the exemplary fan section 14 includes an annular fan casing or outernacelle 50 that circumferentially surrounds the fan 38 and/or at least aportion of the core turbine engine 16. It should be appreciated thatnacelle 50 may be configured to be supported relative to the coreturbine engine 16 by a plurality of circumferentially-spaced outletguide vanes 52. Moreover, a downstream section 54 of the nacelle 50 mayextend over an outer portion of the core turbine engine 16 so as todefine a bypass airflow passage 56 therebetween.

During operation of the turbofan engine 10, a volume of air 58 entersturbofan 10 through an associated inlet 60 of the nacelle 50 and/or fansection 14. As the volume of air 58 passes across fan blades 40, a firstportion of the air 58 as indicated by arrows 62 is directed or routedinto the bypass airflow passage 56 and a second portion of the air 58 asindicated by arrows 64 is directed or routed into the LP compressor 22.The ratio between the first portion of air 62 and the second portion ofair 64 is commonly known as a bypass ratio. The pressure of the secondportion of air 64 is then increased as it is routed through the highpressure (HP) compressor 24 and into the combustion section 26, where itis mixed with fuel and burned to provide combustion gases 66.

The combustion gases 66 are routed through the HP turbine 28 where aportion of thermal and/or kinetic energy from the combustion gases 66 isextracted via sequential stages of HP turbine stator vanes 68 that arecoupled to the outer casing 18 and HP turbine rotor blades 70 that arecoupled to the HP shaft or spool 34, thus causing the HP shaft or spool34 to rotate, thereby supporting operation of the HP compressor 24. Thecombustion gases 66 are then routed through the LP turbine 30 where asecond portion of thermal and kinetic energy is extracted from thecombustion gases 66 via sequential stages of LP turbine stator vanes 72that are coupled to the outer casing 18 and LP turbine rotor blades 74that are coupled to the LP shaft or spool 36, thus causing the LP shaftor spool 36 to rotate, thereby supporting operation of the LP compressor22 and/or rotation of the fan 38.

The combustion gases 66 are subsequently routed through the jet exhaustnozzle section 32 of the core turbine engine 16 to provide propulsivethrust. Simultaneously, the pressure of the first portion of air 62 issubstantially increased as the first portion of air 62 is routed throughthe bypass airflow passage 56 before it is exhausted from a fan nozzleexhaust section 76 of the turbofan 10, also providing propulsive thrust.The HP turbine 28, the LP turbine 30, and the jet exhaust nozzle section32 at least partially define a hot gas path 78 for routing thecombustion gases 66 through the core turbine engine 16.

It will be appreciated that, although described with respect to turbofan10 having core turbine engine 16, the present subject matter may beapplicable to other types of turbomachinery. For example, the presentsubject matter may be suitable for use with or in turboprops,turboshafts, turbojets, industrial and marine gas turbine engines,and/or auxiliary power units.

In some embodiments, components of turbofan engine 10, particularlycomponents within hot gas path 78, such as components of combustionsection 26, HP turbine 28, and/or LP turbine 30, may comprise a ceramicmatrix composite (CMC) material, which is a non-metallic material havinghigh temperature capability. Of course, other components of turbofanengine 10, such as components of HP compressor 24, may comprise a CMCmaterial. Exemplary CMC materials utilized for such components mayinclude silicon carbide (SiC), silicon, silica, or alumina matrixmaterials and combinations thereof. Ceramic fibers may be embeddedwithin the matrix, such as oxidation stable reinforcing fibers includingmonofilaments like sapphire and silicon carbide (e.g., Textron's SCS-6),as well as rovings and yarn including silicon carbide (e.g., NipponCarbon's NICALON®, Ube Industries' TYRANNO®, and Dow Corning'sSYLRAMIC®), alumina silicates (e.g., Nextel's 440 and 480), and choppedwhiskers and fibers (e.g., Nextel's 440 and SAFFIL®), and optionallyceramic particles (e.g., oxides of Si, Al, Zr, Y, and combinationsthereof) and inorganic fillers (e.g., pyrophyllite, wollastonite, mica,talc, kyanite, and montmorillonite). For example, in certainembodiments, bundles of the fibers, which may include a ceramicrefractory material coating, are formed as a reinforced tape, such as aunidirectional reinforced tape. A plurality of the tapes may be laid uptogether (e.g., as plies) to form a preform component. The bundles offibers may be impregnated with a slurry composition prior to forming thepreform or after formation of the preform. The preform may then undergothermal processing, such as a cure or burn-out to yield a high charresidue in the preform, and subsequent chemical processing, such asmelt-infiltration or chemical vapor infiltration with silicon, to arriveat a component formed of a CMC material having a desired chemicalcomposition. In other embodiments, the CMC material may be formed as,e.g., a carbon fiber cloth rather than as a tape.

As stated, components comprising a CMC material may be used within thehot gas path 78, such as within the combustion and/or turbine sectionsof engine 10. As an example, the combustion section 26 may include acombustor formed from a CMC material and/or one or more stages of one ormore stages of the HP turbine 28 may be formed from a CMC material.However, CMC components may be used in other sections as well, such asthe compressor and/or fan sections. Of course, in some embodiments,other high temperature materials and/or other composite materials may beused to form one or more components of engine 10.

FIG. 2 provides an exploded view of a schematic cross-section of thecombustion section 26 and the HP turbine 28 of the turbine section ofthe turbofan engine 10 according to an exemplary embodiment of thepresent subject matter. FIG. 3 provides an unexploded schematiccross-sectional view of the combustion section 26 and the HP turbine 28of FIG. 2 that focuses on an outer boundary of a flow path through thecombustion section 26 and HP turbine 28. The depicted combustion section26 includes a generally annular combustor 80, and downstream of thecombustion section 26, the HP turbine 28 includes a plurality of turbinestages. More particularly, for the depicted embodiment, HP turbine 28includes a first turbine stage 82 and a second turbine stage 84. Inother embodiments, the HP turbine 28 may comprise a different number ofturbine stages; for example, the HP turbine 28 may include one turbinestage or more than two turbine stages. The first turbine stage 82 ispositioned immediately downstream of the combustion section 26, and thesecond turbine stage 84 is positioned immediately downstream of thefirst turbine stage 82. Further, each turbine stage 82, 84 comprises anozzle portion and a blade portion; the first turbine stage 82 includesnozzle portion 82N and blade portion 82B, and the second turbine stage84 includes nozzle portion 84N and blade portion 84B. The nozzle portion82N of the first turbine stage 82 is located immediately downstream ofthe combustion section 26, such that the nozzle portion 82N of the firstturbine stage 82 also may be referred to as a combustor dischargenozzle. Moreover, combustor 80 defines a generally annular combustionchamber 86 such that the combustor 80 may be described as a generallyannular combustor.

Additionally, as described in greater detail below, a flow path 100through the combustion section 26 and the HP turbine 28 is defined by anouter boundary and an inner boundary of a flow path assembly 101. Theouter and inner boundaries form a flow path for the combustion gases 66through the combustion section 26 and HP turbine 28; thus, the flow path100 may comprise at least a portion of the hot gas path 78 describedabove. Further, in other embodiments, the flow path 100 also may extendthrough LP turbine 30 and jet exhaust 32; in still other embodiments,the flow path 100 also may extend forward upstream of the combustionsection 26, e.g., into HP compressor 24. As such, it will be appreciatedthat the discussion herein of the present subject matter with respect tocombustion section 26 and HP turbine 28 is by way of example only andalso may apply to different configurations of gas turbine engines andflow paths 100.

As shown in the exploded view of FIG. 2, the outer and inner boundariesmay be defined by an outer wall 102 and an inner wall 120, respectively,which may include several portions of the combustion section 26 and HPturbine 28. For instance, the combustor 80 includes an outer liner 108defining an outer boundary of the flow path through the combustor 80.Each nozzle portion 82N, 84N comprises an outer band defining an outerboundary of a flow path through the nozzle portion of each turbinestage, and each blade portion 82B, 84B comprises a shroud defining anouter boundary of a flow path through the blade portion of each turbinestage. More particularly, as shown in FIG. 2, the first turbine stagenozzle portion 82N comprises outer band 110, first turbine stage bladeportion 82B comprises shroud 112, second turbine stage nozzle portion84N comprises outer band 114, and second turbine stage blade portion 84Bcomprises shroud 116. These portions of the combustion section 26 and HPturbine 28 may comprise at least a portion of the outer wall 102, asdescribed in greater detail below.

Further, as illustrated in FIG. 2, the combustor 80 includes an innerliner 122 defining an inner boundary of the flow path through thecombustor 80. Each nozzle portion 82N, 84N comprises an inner banddefining an inner boundary of the flow path through the nozzle portionof each turbine stage, and each blade portion 82B, 84B comprises one ormore blade platforms that define an inner boundary of the flow paththrough the blade portion of each turbine stage. More particularly, asshown in FIG. 2, the first turbine stage nozzle portion 82N comprisesinner band 124, first turbine stage blade portion 82B comprises bladeplatforms 132, second turbine stage nozzle portion 84N comprises innerband 136, and second turbine stage blade portion 84B comprises bladeplatforms 132. These portions of the combustion section 26 and HPturbine 28 may comprise at least a portion of the inner wall 122, asdescribed in greater detail below.

Moreover, in the depicted embodiment, a combustor dome 118 extendsradially across a forward end 88 of the combustor 80. The combustor dome118 may be a part of outer wall 102, may be a part of inner wall 120,may be a part of both outer wall 102 and inner wall 120 (e.g., a portionof the combustor dome 118 may be defined by the outer wall 102 and theremainder may be defined by the inner wall 120), or may be a separatecomponent from outer wall 102 and inner wall 120. Additionally, aplurality of nozzle airfoils is positioned in each of the nozzleportions 82N, 84N. Each nozzle airfoil 126 within the first turbinestage nozzle portion 82N extends radially from the outer band 110 to theinner band 124, and the nozzle airfoils 126 are spaced circumferentiallyabout the longitudinal centerline 12. Each nozzle airfoil 128 within thesecond turbine stage nozzle portion 84N extends radially from the outerband 114 to the inner band 136, and the nozzle airfoils 128 are spacedcircumferentially about the longitudinal centerline 12. Further, aplurality of blade airfoils 130 are positioned in each of the bladeportions 82B, 84B. Each blade airfoil 130 within the first turbine stageblade portion 82B is attached to blade platform 132, which in turn isattached to a first stage rotor 134. The blade airfoils 130 attached tothe first stage rotor 134 are spaced circumferentially about thelongitudinal centerline 12. Similarly, each blade airfoil 130 within thesecond turbine stage blade portion 84B is attached to a blade platform132, which in turn is attached to a second stage rotor 138. The bladeairfoils 130 attached to the second stage rotor 138 are spacedcircumferentially about the longitudinal centerline 12. Each bladeairfoils 130 extends radially outward toward the outer wall 102, i.e.,the outer boundary of the flow path 100, and a clearance gap is definedbetween a tip 140 of each blade airfoil 130 and the outer wall 102 suchthat each turbine rotor 134, 138 is free to rotate within its respectiveturbine stage. Although not depicted, each turbine rotor 134, 138 of theHP turbine 28 is connected to the HP shaft 34 (FIG. 1). In such manner,rotor blade airfoils 130 may extract kinetic energy from the flow ofcombustion gases through the flow path 100 defined by the HP turbine 28as rotational energy applied to the HP shaft 34.

Accordingly, flow path 100 through the combustion section 26 and the HPturbine 28 is defined by a flow path assembly 101 having an innerboundary and an outer boundary, and the inner and outer boundariesdefine the flow path for the combustion gases 66 through the combustionsection 26 and HP turbine 28. Portions of the outer boundary of the flowpath assembly 101 may be integrated or unified into a single piece outerwall 102 that defines the radially outer boundary of the gas flow path100. For instance, the outer wall 102 may include a combustor portion104 extending through a combustion section, such as combustion section26, and a turbine portion 106 extending through at least a first turbinestage of a turbine section, such as first turbine stage 82 of HP turbine28. The combustor portion 104 and turbine portion 106 are integrallyformed such that the combustor portion and the turbine portion are asingle unitary structure, i.e., a unitary outer wall 102.

In the exemplary embodiment depicted in FIG. 3A, the outer wall 102includes a combustor portion 104 extending through the combustionsection 26 and a turbine portion 106 extending through at least thefirst turbine stage 82 and the second turbine stage 84 of the turbinesection. In other embodiments, the turbine portion 106 may extendthrough fewer stages (e.g., through one turbine stage as just described)or through more stages (e.g., through one or more stages of the LPturbine 30 positioned downstream of HP turbine 28). The combustorportion 104 and the turbine portion 106 are integrally formed such thatthe combustor portion 104 and the turbine portion 106 are a singleunitary structure, which is referred to herein as unitary outer wall102.

The term “unitary” as used herein denotes that the associated component,such as the outer wall 102, is made as a single piece duringmanufacturing, i.e., the final unitary component is a single piece.Thus, a unitary component has a construction in which the integratedportions are inseparable and is different from a component comprising aplurality of separate component pieces that have been joined togetherand, once joined, are referred to as a single component even though thecomponent pieces remain distinct and the single component is notinseparable (i.e., the pieces may be re-separated). The final unitarycomponent may comprise a substantially continuous piece of material, orin other embodiments, may comprise a plurality of portions that arepermanently bonded to one another. In any event, the various portionsforming a unitary component are integrated with one another such thatthe unitary component is a single piece with inseparable portions.

As shown in FIG. 3, the combustor portion 104 of the unitary structureforming outer wall 102 includes the outer liner 108 of the combustor 80.The turbine portion 106 includes the outer band 110 of the first turbinestage nozzle portion 82N, the shroud 112 of the first turbine stageblade portion 82B, the outer band 114 of the second turbine stage nozzleportion 84N, and the shroud 116 of the second turbine stage bladeportion 84B. As stated, these outer boundary components are integratedinto a single piece to form the unitary structure that is outer wall102. Thus, in the exemplary embodiment of FIG. 2, outer liner 108, outerband 110, shroud 112, outer band 114, and shroud 116 are integrallyformed, i.e., constructed as a single unit or piece to form theintegrated or unitary outer wall 102.

In some embodiments, other portions of the flow path assembly 101 may beintegrated into the unitary structure of outer wall 102, and in stillother embodiments, at least a portion of the outer boundary and theinner boundary are made as a single, unitary component such that theflow path assembly 101 may be referred to as an integrated flow pathassembly. Thus, in various embodiments, the outer wall 102 may includevarious portions of the components along the outer portion of the flowpath 100 and the inner wall 120 may include various portions of thecomponents along the inner portion of the flow path 100. For instance,as shown in FIG. 3, the outer wall 102 may be a unitary outer wall 102where the outer liner 108, outer band 110, shroud 112, outer band 114,and shroud 116 are integrally formed, i.e., constructed as a single unitor piece to form the integrated or unitary outer wall 102. In anotherembodiment, the outer wall 102 may include a portion of the combustordome 118 or may be integrated with the entire combustor dome 118 suchthat the combustor dome 118 and one or more portions of the outersection of the flow path 100 are a single, integral piece. In stillother embodiments, the inner wall 120 may include a portion of thecombustor dome 118 or may be integrated with the entire combustor dome118 such that the combustor dome 118 and one or more portions of theinner section of the flow path 100 are a single, integral piece. Forexample, the flow path assembly 101 may include an outer wall 102 thatcomprises a radially outer portion of the combustor dome 118 and theouter liner 108, which are integrally formed from a CMC material as asingle unit or piece, and an inner wall 120 that comprises a radiallyinner portion of the combustor dome 118 and the inner liner 108, whichare integrally formed from a CMC material as a single unit or piece.

In yet other embodiments, the combustor dome 118 may not be integratedwith either the outer wall 102 or the inner wall 120 in whole or inpart. That is, the combustor dome 118 is a separate component from boththe outer wall 102 and the inner wall 120. As such, the flow path 100may be discontinuous between the combustor dome 118 and outer wall 102,as well as between combustor dome 118 and inner wall 120. Further, insuch embodiments, the combustor dome 118 is configured to move axiallywith respect to the inner wall 120 and the outer wall 102 but may beattached to, and accordingly supported by, one or more fuel nozzleassemblies 90. More particularly, an axial slip joint may be formedbetween the combustor dome 118 and each of the outer wall 102 and theinner wall 120 such that the combustor dome 118 may move or floataxially with respect to the inner wall 120 and outer wall 102. Allowingthe combustor dome 118 to float relative to the outer wall 102 and innerwall 120 can help control the position of the fuel nozzle assembly 90with respect to the combustor dome 118 and combustor 80. For example,the combustor dome 118, outer wall 102, and inner wall 120 may be madeof a different material or materials than the fuel nozzle assembly 90.As described in greater detail below, in an exemplary embodiment, thecombustor dome 118, outer wall 102, and inner wall 120 are made from aceramic matrix composite (CMC) material, and the fuel nozzle assembly 90may be made from a metallic material, e.g., a metal alloy or the like.In such embodiment, the CMC material thermally grows or expands at adifferent rate than the metallic material. Thus, allowing the combustordome 118 to move axially with respect to outer and inner walls 102, 120may allow for tighter control of the immersion of swirler 92 of fuelnozzle assembly 90 within combustor dome 118, as well as combustor 80,than if the combustor dome 118 was attached to the outer and inner walls102, 120. Tighter control of the position of fuel nozzle assembly 90 andits components with respect to combustor 80 can reduce variation inoperability and performance of engine 10.

Additionally, in embodiments in which the combustor dome 118 is separatefrom the outer and inner walls 102, 120, the outer wall 102 and innerwall 120 also may move axially and radially with respect to thecombustor dome 118. By decoupling the combustor dome 118 from the walls102, 120 and allowing relative movement between the walls 102, 120 andthe combustor dome 118, stress coupling may be alleviated between theouter and inner walls 102, 120 and the combustor dome 118. Moreover, anyleakage between the uncoupled combustor dome 118 and outer and innerwalls 102, 120 may be utilized as purge and/or film starter flow.

FIG. 4 provides a partial perspective view of a portion of an integralflow path assembly 101, having an outer wall 102 and inner wall 120formed as a single piece component. As shown in FIG. 4, in someembodiments of the combustion gas flow path assembly 101, the outerliner 108, outer band 110, shroud 112, outer band 114, shroud 116,combustor dome 118, inner liner 122, and inner band 124 are integrallyformed such that the outer liner 108, outer bands 110, 114, shrouds 112,116, combustor dome 118, inner liner 122, and inner band 124 are asingle unitary structure. FIG. 4 further illustrates that a plurality ofopenings 142 for receipt of fuel nozzle assemblies 90 and/or swirlers 92may be defined in the forward end 88 of combustor 80 of the unitary flowpath assembly 101. Further, it will be appreciated that FIG. 4illustrates only a portion of the integral flow path assembly 101 andthat, although its entire circumference is not illustrated in FIG. 4,the flow path assembly 101 is a single, unitary piece circumferentiallyas well as axially. As such, the integral flow path assembly 101 definesa generally annular, i.e., generally ring-shaped, flow path between theouter wall 102 and inner wall 120.

Integrating various components of the outer and inner boundaries of theflow path assembly 101 as described above can reduce the number ofseparate pieces or components within engine 10, as well as reduce theweight, leakage, and complexity of the engine 10, compared to known gasturbine engines. For instance, known gas turbine engines employ seals orsealing mechanisms at the interfaces between separate pieces of the flowpath assembly to attempt to minimize leakage of combustion gases fromthe flow path. By integrating the outer boundary, for example, asdescribed with respect to unitary outer wall 102, split points orinterfaces between the outer combustor liner and first turbine stageouter band, the first turbine stage outer band and the first turbinestage shroud, etc. can be eliminated, thereby eliminating leakage pointsas well as seals or sealing mechanisms required to prevent leakage.Similarly, by integrating components of the inner boundary, split pointsor interfaces between the integrated inner boundary components areeliminated, thereby eliminating leakage points and seals or sealingmechanisms required at the inner boundary. Accordingly, undesiredleakage, as well as unnecessary weight and complexity, can be avoided byutilizing unitary components in the flow path assembly. Other advantagesof unitary outer wall 102, unitary inner wall 120, and/or a unitary flowpath assembly 101 will be appreciated by those of ordinary skill in theart.

As most clearly illustrated in FIG. 4, the outer wall 102 and the innerwall 120 define a generally annular flow path therebetween. That is, theunitary outer wall 102 circumferentially surrounds the inner wall 120;stated differently, the unitary outer wall 102 is a single pieceextending 360° degrees about the inner wall 120, thereby defining agenerally annular or ring-shaped flow path therebetween. As such, thecombustor dome 118, which extends across the forward end 88 of thecombustor 80, is a generally annular combustor dome 118. Further, thecombustor dome 118 defines an opening 142 for receipt of a fuel nozzleassembly 90 positioned at forward end 88. The fuel nozzle assembly 90,e.g., provides combustion chamber 86 with a mixture of fuel andcompressed air from the compressor section, which is combusted withinthe combustion chamber 86 to generate a flow of combustion gases throughthe flow path 100. The fuel nozzle assembly 90 may attach to thecombustor dome 118 or may “float” relative to the combustor dome 118 andthe flow path 100, i.e., the fuel nozzle assembly 90 may not be attachedto the combustor dome 118. In the illustrated embodiments, the fuelnozzle assembly 90 includes a swirler 92, and in some embodiments, theswirler 92 may attach to the combustor dome 118, but alternatively, theswirler 92 may float relative to the combustor dome 118 and flow path100. It will be appreciated that the fuel nozzle assembly 90 or swirler92 may float relative to the combustor dome 118 and flow path 100 alongboth a radial direction R and an axial direction A or only along one orthe other of the radial and axial directions R, A. Further, it will beunderstood that the combustor dome 118 may define a plurality ofopenings 142, each opening receiving a swirler 92 or other portion offuel nozzle assembly 90.

As further illustrated in FIGS. 2, 3, and 4, the flow path assembly 101generally defines a converging-diverging flow path 100. Moreparticularly, the outer wall 102 and the inner wall 120 define agenerally annular combustion chamber 86, which forms a forward portionof the flow path 100. Moving aft or downstream of combustion chamber 86,the outer wall 102 and inner wall 120 converge toward one another,generally in the region of first turbine stage 82. Continuing downstreamof the first turbine stage 82, the outer wall 102 and inner wall 120then diverge, generally in the region of second turbine stage 84. Theouter wall 102 and inner wall 120 may continue to diverge downstream ofthe second turbine stage 84. In exemplary embodiments, e.g., as shown inFIG. 3 and referring only to the unitary outer wall 102, the firstturbine stage nozzle outer band portion 110 and blade shroud portion 112of the outer wall 102 converge toward the axial centerline 12. Thesecond turbine stage nozzle outer band portion 114 and blade shroudportion 116 of the outer wall 102 diverge away from the axial centerline12. As such, the outer boundary of flow path 100 formed by the unitaryouter wall 102 defines a converging-diverging flow path 100.

Turning to FIG. 5, a perspective view is provided of a nozzle airfoil ofthe flow path assembly 101. As shown in the depicted embodiment, theflow path assembly 101 includes an inner wall 120 and a unitary outerwall 102. As described above, the unitary outer wall 102 includes acombustor portion 104 that extends through the combustion section 26 anda turbine portion 106 that extends through at least a first turbinestage 82 of the turbine section 28; in the embodiment of FIG. 5, theturbine portion extends through the second turbine stage 84. Further,the combustor portion 104 and the turbine portion 106 of the outer wall102 are integrally formed as a single unitary structure and, thus, maybe referred to as unitary outer wall 102, and the inner wall 120 and theunitary outer wall 102 define the combustor 80.

More particularly, in the illustrated embodiment, the combustor portion104 of the unitary outer wall 102 comprises the outer liner 108 of thecombustor 80, and the turbine portion 106 comprises the outer band 110of the first turbine stage nozzle portion 82N, the shroud 112 of thefirst turbine stage blade portion 82B, the outer band 114 of the secondturbine stage nozzle portion 84N, and the shroud 116 of the secondturbine stage blade portion 84B. The inner wall 120 also may be aunitary structure that may be referred to as unitary inner wall 120; forexample, in FIG. 5, the inner wall 120 is a unitary structure comprisingthe inner liner 122 and first turbine stage inner band 124, which areintegrally formed as unitary inner wall 120. In some embodiments, aspreviously described, the unitary outer wall 102 or unitary inner wall120 also may include the combustor dome 118, or the unitary outer wall102 and the unitary inner wall 120 each may include a portion of thecombustor dome 118. In still other embodiments, the outer wall 102,combustor dome 118, and inner wall 120 may be integrally formed as asingle piece, unitary structure.

Referring still to the embodiment illustrated in FIG. 5, the flow pathassembly 101 further includes a plurality of first turbine stage nozzleairfoils 126. Each of the plurality of first turbine stage nozzleairfoils 126 has an inner end 126 a that is radially opposite an outerend 126 b. Further, each of the plurality of nozzle airfoils 126 extendsinto the inner wall 120 and radially outward through the outer wall 102.More specifically, the outer wall 102 defines a plurality of openings170 therethrough, which are circumferentially spaced apart from oneanother. Each opening 170 is configured for receipt of one of theplurality of first turbine stage nozzle airfoils 126, such that theouter end 126 b of each first turbine stage nozzle airfoil 126 extendsthrough an opening 170. Accordingly, each opening 170 preferably has ashape substantially similar to an axial cross-sectional shape of thefirst turbine stage nozzle airfoils 126. That is, each opening 170 maygenerally be described as an airfoil-shaped cutout in the outer wall102. However, the openings 170 may have any appropriate shape forreceiving the nozzle airfoils 126.

Further, the inner wall 120 defines a plurality of pockets 174, whichare circumferentially spaced apart from one another. Each of the pockets174 is configured to receive one of the plurality of first turbine stagenozzle airfoils 126. That is, the inner end 126 a of each first turbinestage nozzle airfoil 126 is received within a pocket 174. Similar to theopenings 170, the pockets 174 may have any suitable shape for receipt ofthe nozzle airfoils 126, e.g., each pocket 174 may be substantiallyairfoil shaped, but other suitable shapes may be used as well.

As further illustrated in the exemplary embodiment of FIG. 5, eachnozzle airfoil 126 has a concave pressure side 180 opposite a convexsuction side 182. Opposite pressure and suction sides 180, 182 of eachnozzle airfoil 126 radially extend between the inner end 126 a and theouter end 126 b. Moreover, pressure and suction sides 180, 182 of nozzleairfoils 126 axially extend between a leading edge 184 and an oppositetrailing edge 186. Leading edge 184 defines a forward end of nozzleairfoil 126, and trailing edge 186 defines an aft end of nozzle airfoil126. Further, each nozzle airfoil 126 defines a chord c extendingaxially between the opposite leading and trailing edges 184, 186.Pressure and suction sides 180, 182 of each nozzle airfoil 126 define anouter surface 188 of the airfoil. Additionally, each nozzle airfoil 126may define one or move internal cavities 190 for receiving a flow ofcooling fluid F, e.g., a flow of pressurized air diverted from HPcompressor 24. More particularly, a fluid flowing through the flow path100, such as the combustion gases 66, may have a flow path pressure, andthe flow of cooling fluid F may be pressurized above the flow pathpressure. Further, each internal cavity 190, in turn, may providecooling to one or more portions of airfoil 126, as well as to pockets174 as described in greater detail below.

Referring still to FIG. 5, each nozzle airfoil 126 defines an inlet 191for receipt of the flow of cooling fluid F into the internal cavity 190,as well as an outlet 192 for the flow of cooling fluid F to flow fromthe internal cavity 190 to the pocket 174. As such, the outlet 192 isdefined at the inner end 126 a of each nozzle airfoil 126. Moreover, theinlet 191 or the outlet 192 of each nozzle airfoil 126 may be sized tometer the flow of cooling fluid F from the internal cavity 190 to thepocket 174. For instance, the inlet 191 or the outlet 192 may have adiameter, cross-sectional area, or other size parameter that is selectedto provide a cooling fluid flow within a certain range of fluid flowrates to the respective pocket 174, e.g., under a specific operatingcondition or a range of operating conditions. Thus, the cooling fluidflow rate may be a controlled flow rate such that the pockets 174 arepurged at a controlled purge flow rate.

Further, at least a substantial portion of the cooling fluid F flowingfrom the outlet 192 of each nozzle airfoil 126 is directed by thepockets 174 such that the cooling fluid F flows up along the pressureand suction sides 180, 182 of the nozzle airfoils 126. The flow ofcooling fluid F along the outer surface 188 of the nozzle airfoils 126forms a fluid curtain C across the airfoils 126 that, e.g., discouragesfluid leakage from the flow path 100. More particularly, the fluidcurtain C formed by the flow of cooling fluid F from the internal cavity190 of a nozzle airfoil 126 discourages cross-over leakage from thepressure side 180 to the suction side 182, e.g., by blocking the fluidthat otherwise may cross-over from the pressure side 180 to the suctionside 182 via pocket 174.

In other embodiments, apertures may be defined through inner wall 120 toprovide a flow of cooling fluid to pockets 174. For example, as shown inFIG. 5, one or more apertures 194 may be defined through the inner wall120 to provide a fluid flow to a pocket 174, i.e., each aperture 194 isdefined from an outer surface 121 of the inner wall 120 to the pocket174 such that the cooling fluid F can flow from outside the flow pathassembly 101 to the pocket 174. Similar to the flow of cooling fluid Ffrom internal cavities 190, at least a substantial portion of thecooling fluid F flowing from the aperture(s) 194 of a pocket 174 isdirected by the pocket 174 such that the cooling fluid F flows up alongthe pressure and suction sides 180, 182 of the nozzle airfoils 126. Theflow of cooling fluid F thereby forms a fluid curtain C across the outersurface 188 of the airfoils 126 that, e.g., discourages fluid leakage,such as cross-over pressure side to suction side leakage, from the flowpath 100. Further, the apertures 194 may be sized to meter to the flowrate of cooling fluid F from the apertures, i.e., the diameter,cross-sectional area, or the like of each aperture 194 may be selectedto provide a certain flow rate or range of flow rates under a givenoperating condition or range of operating conditions. Thus, like coolingfluid from internal cavities 190, the flow of cooling fluid F fromapertures 194 may provide a controlled purge flow rate to pockets 174.

It will be appreciated that the apertures 194 may be provided inaddition to or as an alternative to cavity 190. That is, a flow ofcooling fluid F may be provided to pockets 174 from both cavities 190and apertures 194, from cavities 190 only, or from apertures 194 only.In some embodiments, each nozzle airfoil 126 may define an internalcavity 190 that provides a flow of cooling fluid F through an outlet 192in the nozzle airfoil 126, and a portion of the plurality of pockets 174also may be provided with a flow of cooling fluid F from one or moreapertures 194 defined in the inner wall 120. As such, the combination offluid flow from an internal cavity 190 and one or more apertures 194 maybe utilized for some but not all of the nozzle airfoils 126 and pockets174 of flow path assembly 101. For example, the flow of cooling fluid Ffrom internal cavity 190 may be supplemented by a flow of cooling fluidF from apertures 194 in certain areas of the flow path assembly 101 thatmay experience increased temperatures, i.e., in hot spots of the flowpath.

The flow of cooling fluid F from the internal cavities 190 or otherapertures may have other uses as well. As an example, the cooling fluidF may help cool the pockets 174 and downstream airfoil root fillets.Additionally, the flow of cooling fluid F within internal cavities 190may help cool the nozzle airfoils 126, and the internal cavity 190 ofeach nozzle airfoil 126 may be defined in an area that has the greatestcooling need, e.g., some areas of airfoils 126 may require cooling morethan other areas, and as such, the internal cavities 190 may be definedin airfoils 126 to provide cooling to such areas. Moreover, the flow ofcooling fluid F to pockets 174 may eliminate the need for separate sealsbetween the inner wall 120 and inner end 126 a of nozzle airfoils 126,e.g., the cooling fluid may provide a fluid seal between the inner wall120 and the airfoils 126. However, in some embodiments, the flow pathassembly 101 includes a plurality of seals, such as wire seals or othersuitable seals, and a seal also may be provided within each pocket 174between inner wall 120 and the inner end 126 a of nozzle airfoils 126.

Further, the nozzle airfoils 126 may be secured in the openings 170and/or pockets 174 using any suitable means. For instance, one or moreretention mechanisms, such as pins, loading rings, or the like, may beused to secure the nozzle airfoils 126 against retracting from theopenings 170, 172. As another example, the nozzle airfoils 126 may bejoined to the outer wall 102 to help secure and retain the airfoilswithin the openings 170. In one embodiment, each nozzle airfoil 126 maybe bonded to the outer wall 102, e.g., at the opening 170 in which therespective airfoil 126 is received. In other embodiments, one or moreseals may be positioned at the interface between the nozzle airfoils 126and the outer wall 102 to prevent hot gas leakage through openings 170.

It will be appreciated that, although FIG. 5 depicts only a firstturbine stage nozzle airfoil 126 of the flow path assembly 101, thesubject matter described with respect to the first turbine stage nozzleairfoils 126 applies equally to the second turbine stage nozzle airfoils128, and also may apply to any other nozzle portions of the turbinesections of the engine 10. That is, while only the first turbine stagenozzle airfoils 126 are shown and described with respect to FIG. 5, thesecond turbine stage nozzle airfoils 128 may be similarly configured,and the second stage inner band 136 may be configured similarly to thefirst stage inner band portion of the inner wall 120, e.g., with pocketsdefined therein for receipt of the inner end of nozzle airfoils 128.Further, in some embodiments, apertures 194 also may be defined throughthe second stage inner band 136 to provide a flow of cooling fluid F tothe pockets defined therein. However, in other embodiments, the firstturbine stage nozzle portion 82N may be configured as shown in FIG. 5while the second turbine stage nozzle portion 84N, as well as nozzleportions of other turbine stages, may be configured differently from thedepicted embodiment.

Also, it will be appreciated that the embodiment shown in FIG. 5 is byway of example only, and other configurations of flow path assembly 101utilizing drop-in nozzle airfoils may be used as well. For example, inone embodiment, the configuration shown in FIG. 5 may be reversed, i.e.,the openings 170 may be defined in the inner wall 120 and the pockets174 may be defined in the outer wall 102, such that the nozzle airfoils126 are inserted through the inner boundary structure of flow pathassembly 101 and into the outer boundary structure, where the outer end126 b of each airfoil 126 is received in an outer wall pocket 174. Theinner cavities 190 of nozzle airfoils 126 may then provide a flow ofcooling fluid F through outlets 192 defined in the outer ends 126 b topurge the pockets 174 defined in the outer wall 102 and to form a fluidcurtain C along the outer surface 188 of the airfoils 126. Further, insuch embodiments, apertures 194 may be defined in the outer wall 102,e.g., from an outer surface of the outer wall 102 to the outer wallpockets 174, to provide a flow of cooling fluid F through the apertures194, in addition to or in place of the flow of cooling fluid F providedthrough the internal cavities 190. Of course, other configurations ofthe flow path assembly 101 also may be used.

As previously stated, the outer wall 102, inner wall 120, and combustordome 118, and in some embodiments, first and second turbine stage nozzleairfoils 126, 128, may comprise a CMC material. More particularly, inexemplary embodiments, the combustor portion 104 and the turbine portion106 of flow path assembly 101 are integrally formed from a CMC materialsuch that the resulting unitary structure is a CMC component. Forexample, where the combustor portion 104 includes the outer liner 108 ofthe combustor 80 and the turbine portion 106 includes the outer band 110of the first turbine stage nozzle portion 82N, the shroud 112 of thefirst turbine stage blade portion 82B, the outer band 114 of the secondturbine stage nozzle portion 84N, and the shroud 116 of the secondturbine stage blade portion 84B, the outer liner 108, outer bands 110,114, and shrouds 114, 116 may be integrally formed from a CMC materialto produce a unitary CMC outer wall 102. As described above, in otherembodiments, additional CMC components may be integrally formed with theouter liner 108, outer bands 110, 114, and shrouds 114, 116 to constructa unitary CMC outer wall 102. Similarly, the inner wall 120 may beformed from a CMC material. For instance, where the inner wall 120comprises separate components, e.g., inner liner 122, inner bands 124,136, and blade platforms 132, each component of the inner wall 120 maybe formed from a CMC material. In embodiments in which two or morecomponents are integrated to form a unitary inner wall 120, thecomponents may be integrally formed from a CMC material to construct aunitary CMC inner wall 120.

Examples of CMC materials, and particularly SiC/Si—SiC (fiber/matrix)continuous fiber-reinforced ceramic composite (CFCC) materials andprocesses, are described in U.S. Pat. Nos. 5,015,540; 5,330,854;5,336,350; 5,628,938; 6,024,898; 6,258,737; 6,403,158; and 6,503,441,and U.S. Patent Application Publication No. 2004/0067316. Such processesgenerally entail the fabrication of CMCs using multiple pre-impregnated(prepreg) layers, e.g., the ply material may include prepreg materialconsisting of ceramic fibers, woven or braided ceramic fiber cloth, orstacked ceramic fiber tows that has been impregnated with matrixmaterial. In some embodiments, each prepreg layer is in the form of a“tape” comprising the desired ceramic fiber reinforcement material, oneor more precursors of the CMC matrix material, and organic resinbinders. Prepreg tapes can be formed by impregnating the reinforcementmaterial with a slurry that contains the ceramic precursor(s) andbinders. Preferred materials for the precursor will depend on theparticular composition desired for the ceramic matrix of the CMCcomponent, for example, SiC powder and/or one or more carbon-containingmaterials if the desired matrix material is SiC. Notablecarbon-containing materials include carbon black, phenolic resins, andfuranic resins, including furfuryl alcohol (C₄H₃OCH₂OH). Other typicalslurry ingredients include organic binders (for example, polyvinylbutyral (PVB)) that promote the flexibility of prepreg tapes, andsolvents for the binders (for example, toluene and/or methyl isobutylketone (MIBK)) that promote the fluidity of the slurry to enableimpregnation of the fiber reinforcement material. The slurry may furthercontain one or more particulate fillers intended to be present in theceramic matrix of the CMC component, for example, silicon and/or SiCpowders in the case of a Si—SiC matrix. Chopped fibers or whiskers orother materials also may be embedded within the matrix as previouslydescribed. Other compositions and processes for producing compositearticles, and more specifically, other slurry and prepreg tapecompositions, may be used as well, such as, e.g., the processes andcompositions described in U.S. Patent Application Publication No.2013/0157037.

The resulting prepreg tape may be laid-up with other tapes, such that aCMC component formed from the tape comprises multiple laminae, eachlamina derived from an individual prepreg tape. Each lamina contains aceramic fiber reinforcement material encased in a ceramic matrix formed,wholly or in part, by conversion of a ceramic matrix precursor, e.g.,during firing and densification cycles as described more fully below. Insome embodiments, the reinforcement material is in the form ofunidirectional arrays of tows, each tow containing continuous fibers orfilaments. Alternatives to unidirectional arrays of tows may be used aswell. Further, suitable fiber diameters, tow diameters, andcenter-to-center tow spacing will depend on the particular application,the thicknesses of the particular lamina and the tape from which it wasformed, and other factors. As described above, other prepreg materialsor non-prepreg materials may be used as well.

After laying up the tapes or plies to form a layup, the layup isdebulked and, if appropriate, cured while subjected to elevatedpressures and temperatures to produce a preform. The preform is thenheated (fired) in a vacuum or inert atmosphere to decompose the binders,remove the solvents, and convert the precursor to the desired ceramicmatrix material. Due to decomposition of the binders, the result is aporous CMC body that may undergo densification, e.g., melt infiltration(MI), to fill the porosity and yield the CMC component. Specificprocessing techniques and parameters for the above process will dependon the particular composition of the materials. For example, silicon CMCcomponents may be formed from fibrous material that is infiltrated withmolten silicon, e.g., through a process typically referred to as theSilcomp process. Another technique of manufacturing CMC components isthe method known as the slurry cast melt infiltration (MI) process. Inone method of manufacturing using the slurry cast MI method, CMCs areproduced by initially providing plies of balanced two-dimensional (2D)woven cloth comprising silicon carbide (SiC)-containing fibers, havingtwo weave directions at substantially 90° angles to each other, withsubstantially the same number of fibers running in both directions ofthe weave. The term “silicon carbide-containing fiber” refers to a fiberhaving a composition that includes silicon carbide, and preferably issubstantially silicon carbide. For instance, the fiber may have asilicon carbide core surrounded with carbon, or in the reverse, thefiber may have a carbon core surrounded by or encapsulated with siliconcarbide.

Other techniques for forming CMC components include polymer infiltrationand pyrolysis (PIP) and oxide/oxide processes. In PIP processes, siliconcarbide fiber preforms are infiltrated with a preceramic polymer, suchas polysilazane and then heat treated to form a SiC matrix. Inoxide/oxide processing, aluminum or alumino-silicate fibers may bepre-impregnated and then laminated into a preselected geometry.Components may also be fabricated from a carbon fiber reinforced siliconcarbide matrix (C/SiC) CMC. The C/SiC processing includes a carbonfibrous preform laid up on a tool in the preselected geometry. Asutilized in the slurry cast method for SiC/SiC, the tool is made up ofgraphite material. The fibrous preform is supported by the toolingduring a chemical vapor infiltration process at about 1200° C., wherebythe C/SiC CMC component is formed. In still other embodiments, 2D, 2.5D,and/or 3D preforms may be utilized in MI, CVI, PIP, or other processes.For example, cut layers of 2D woven fabrics may be stacked inalternating weave directions as described above, or filaments may bewound or braided and combined with 3D weaving, stitching, or needling toform 2.5D or 3D preforms having multiaxial fiber architectures. Otherways of forming 2.5D or 3D preforms, e.g., using other weaving orbraiding methods or utilizing 2D fabrics, may be used as well.

Thus, a variety of processes may be used to form a unitary structure,such as the outer wall 102 depicted in FIG. 3A, as a unitary CMCcomponent. More specifically, a plurality of plies of a CMC material maybe used to form each unitary structure. The plurality of plies may beinterspersed with one another to integrate the various portions formingthe unitary structure. As an example, the unitary outer wall 102 of FIG.3A may be made from a plurality of outer liner plies, a plurality offirst turbine stage outer band plies, a plurality of first turbine stageshroud plies, a plurality of second turbine stage outer band plies, anda plurality of second turbine stage shroud plies. Where the outer linerplies meet the first turbine stage outer band plies, ends of the outerliner plies may be alternated with ends of the outer band plies tointegrate the plies for forming the outer liner portion with the pliesfor forming the first turbine stage outer band portion of the unitaryouter wall 102. That is, any joints between the plies forming unitaryouter wall 102 may be formed by alternating plies on one side of thejoint with plies on the other side of the joint. As such, the plies forforming unitary outer wall 102 may be interspersed to integrate theplies and, thereby, each portion of the unitary outer wall 102. Ofcourse, the CMC plies may be laid up in other ways as well to form theunitary structure. In addition, laying up the plurality of CMC plies mayinclude defining features of the unitary structure or other component(e.g., inner liner 122 when not integrated with inner band 124 to form aunitary inner wall 120 or separate combustor dome 118 as shown in theembodiments of FIGS. 5A and 5B) such as openings 142 in combustorforward end 88, outer wall flange 144, inner wall flange 146, and firstand second openings 170, 172.

After the plurality of CMC plies are laid up to define a unitary CMCcomponent preform, the preform is cured to produce a single piece,unitary CMC component, which is then fired and subjected todensification, e.g., silicon melt-infiltration, to form a final unitaryCMC structure. Continuing with the above outer wall 102 example, theouter wall preform may be processed in an autoclave to produce a greenstate unitary outer wall 102. Then, the green state unitary outer wall102 may be placed in a furnace to burn out excess binders or the likeand then placed in a furnace with a piece or slab of silicon and firedto melt infiltrate the unitary outer wall 102 with at least silicon.More particularly, for unitary outer wall 102 formed from CMC plies ofprepreg tapes that are produced as described above, heating (i.e.,firing) the green state component in a vacuum or inert atmospheredecomposes the binders, removes the solvents, and converts the precursorto the desired ceramic matrix material. The decomposition of the bindersresults in a porous CMC body; the body may undergo densification, e.g.,melt infiltration (MI), to fill the porosity. In the foregoing examplewhere the green state unitary outer wall 102 is fired with silicon, theouter wall 102 undergoes silicon melt-infiltration. However,densification may be performed using any known densification techniqueincluding, but not limited to, Silcomp, melt infiltration (MI), chemicalvapor infiltration (CVI), polymer infiltration and pyrolysis (PIP), andoxide/oxide processes, and with any suitable materials including but notlimited to silicon. In one embodiment, densification and firing may beconducted in a vacuum furnace or an inert atmosphere having anestablished atmosphere at temperatures above 1200° C. to allow siliconor other appropriate material or combination of materials tomelt-infiltrate into the component. The densified CMC body hardens to afinal unitary CMC outer wall 102. In some embodiments, the final unitarystructure may be finish machined, e.g., to bring the structure withintolerance or to define openings 142 in forward end 88 or to defineopenings 170, 172 in the nozzle portions of the turbine stages 82, 84,and/or an environmental barrier coating (EBC) may be applied to theunitary structure, e.g., to protect the unitary structure from the hotcombustion gases 66. It will be appreciated that other methods orprocesses of forming CMC components, such as unitary CMC outer wall 102,unitary CMC inner wall 120, or the like may be used as well.

Additionally or alternatively, other processes for producing unitarycomponents may be used to form unitary outer wall 102 and/or unitaryinner wall 120, and the unitary structure(s) may be formed from othermaterials. In some embodiments, an additive manufacturing process may beused to form unitary outer wall 102 and/or unitary inner wall 120. Forexample, an additive process such as Fused Deposition Modeling (FDM),Selective Laser Sintering (SLS), Stereolithography (SLA), Digital LightProcessing (DLP), Direct Metal Laser Sintering (DMLS), Laser Net ShapeManufacturing (LNSM), electron beam sintering or other known process maybe used to produce a unitary outer wall 102 and/or a unitary inner wall120. Generally, an additive process fabricates components usingthree-dimensional information, for example, a three-dimensional computermodel, of the component. The three-dimensional information is convertedinto a plurality of slices, each slice defining a cross section of thecomponent for a predetermined height of the slice. The component is then“built-up” slice by slice, or layer by layer, until finished. Superalloymetallic materials or other suitable materials may be used in anadditive process to form unitary outer wall 102 and/or a unitary innerwall 120. In other embodiments, a unitary outer wall 102 and/or unitaryinner wall 120 may be formed using a forging or casting process. Othersuitable processes or methods may be used as well.

This written description uses examples to disclose the invention,including the best mode, and also to enable any person skilled in theart to practice the invention, including making and using any devices orsystems and performing any incorporated methods. The patentable scope ofthe invention is defined by the claims and may include other examplesthat occur to those skilled in the art. Such other examples are intendedto be within the scope of the claims if they include structural elementsthat do not differ from the literal language of the claims or if theyinclude equivalent structural elements with insubstantial differencesfrom the literal language of the claims.

What is claimed is:
 1. A flow path assembly for a gas turbine engine,the flow path assembly comprising: an inner wall defining an innerboundary of a flow path, the inner wall further defining a plurality ofpockets therein; an outer wall defining an outer boundary of the flowpath; and a plurality of nozzle airfoils, each nozzle airfoil having aninner end radially opposite an outer end and a pressure side opposite asuction side such that the pressure side and the suction side radiallyextend between the inner end and the outer end, each nozzle airfoildefining an internal cavity for receipt of a flow of cooling fluid,wherein the inner end of each nozzle airfoil is received in a respectivepocket of the plurality of pockets defined in the inner wall, whereinthe outer end of each nozzle airfoil defines an inlet for the flow ofcooling fluid to flow into the internal cavity, wherein the inner end ofeach nozzle airfoil defines an outlet for the flow of cooling fluid toflow from the internal cavity to the respective pocket of the pluralityof pockets, and wherein each pocket of the plurality of pockets isconfigured to direct the cooling fluid flowing from the outlet of eachnozzle airfoil along the pressure side and the suction side of therespective nozzle airfoil to form a fluid curtain to discourage fluidleakage from the flow path.
 2. The flow path assembly of claim 1,wherein the outer wall defines a plurality of openings therethrough, andwherein each opening is configured for receipt of one of the pluralityof nozzle airfoils.
 3. The flow path assembly of claim 2, wherein theouter end of each nozzle airfoil is positioned in one opening of theplurality of openings in the outer wall when the inner end of the nozzleairfoil is received in the respective pocket of the plurality of pocketsdefined in the inner wall.
 4. The flow path assembly of claim 1, furthercomprising a plurality of seals, wherein one seal of the plurality ofseals is positioned in each of the plurality of pockets between theinner end of the nozzle airfoil received in the respective pocket of theplurality of pockets and the inner wall.
 5. The flow path assembly ofclaim 1, wherein a fluid flows through the flow path at a flow pathpressure, and wherein the flow of cooling fluid is pressurized above theflow path pressure.
 6. The flow path assembly of claim 1, wherein theinlet or the outlet is sized to meter the flow of cooling fluid from theinternal cavity to the respective pocket of the plurality of pockets. 7.The flow path assembly of claim 1, wherein the outer wall includes acombustor portion extending through a combustion section of the gasturbine engine and a turbine portion extending through at least a firstturbine stage of a turbine section of the gas turbine engine, whereinthe combustor portion and the turbine portion are integrally formed as asingle unitary structure from a ceramic matric composite (CMC) materialsuch that the outer wall is a unitary CMC component, and wherein theinner wall is formed from a CMC material such that the inner wall is aCMC component.
 8. A flow path assembly for a gas turbine engine, theflow path assembly comprising: an inner wall defining an inner boundaryof a flow path, the inner wall further defining a plurality of pocketstherein; an outer wall defining an outer boundary of the flow path; anda plurality of nozzle airfoils, each nozzle airfoil having an inner endradially opposite an outer end and a pressure side opposite a suctionside such that the pressure side and the suction side radially extendbetween the inner end and the outer end, the inner end of each nozzleairfoil received in a respective pocket of the plurality of pocketsdefined in the inner wall, wherein a flow of cooling fluid is directedinto each pocket of the plurality of pockets, each pocket of theplurality of pockets configured to direct the flow of cooling fluidalong the pressure side of a respective nozzle airfoil of the pluralityof nozzle airfoils received in the respective pocket of the plurality ofpockets to form a fluid curtain to discourage fluid leakage from theflow path into the respective pocket of the plurality of pockets.
 9. Theflow path assembly of claim 8, wherein each nozzle airfoil defines aninternal cavity for receipt of the flow of cooling fluid to direct intoeach pocket of the plurality of pockets.
 10. The flow path assembly ofclaim 9, wherein the inner end of each nozzle airfoil defines an outletto direct the flow of cooling fluid from the internal cavity to therespective pocket of the plurality of pockets.
 11. The flow pathassembly of claim 8, wherein a plurality of apertures are defined in theinner wall, each aperture defined from an outer surface of the innerwall to one of the plurality of pockets.
 12. The flow path assembly ofclaim 11, wherein the flow of cooling fluid is directed through eachaperture to the plurality of pockets.
 13. The flow path assembly ofclaim 8, wherein the fluid curtain is formed along an outer surface ofeach nozzle airfoil.
 14. The flow path assembly of claim 8, wherein aplurality of openings are defined in the outer wall, and wherein theplurality of nozzle airfoils are inserted through the outer wall andinto the pockets in the inner wall such that the outer end of eachnozzle airfoil is positioned in one of the plurality of openings.
 15. Aflow path assembly for a gas turbine engine, the flow path assemblycomprising: an inner wall defining an inner boundary of a flow path; anouter wall defining an outer boundary of the flow path, the outer wallfurther defining a plurality of pockets therein; and a plurality ofnozzle airfoils, each nozzle airfoil having an inner end radiallyopposite an outer end and a pressure side opposite a suction side suchthat the pressure side and the suction side radially extend between theinner end and the outer end, the outer end of each nozzle airfoilreceived in a respective pocket of the plurality of pockets defined inthe outer wall, wherein a flow of cooling fluid is directed into eachpocket of the plurality of pockets, each pocket of the plurality ofpockets configured to direct the flow of cooling fluid along thepressure side of a respective nozzle airfoil of the plurality of nozzleairfoils received in the respective pocket to form a fluid curtain todiscourage fluid leakage from the flow path into the respective pocketof the plurality of pockets.
 16. The flow path assembly of claim 15,wherein each nozzle airfoil defines an internal cavity for receipt ofthe flow of cooling fluid to direct into each pocket of the plurality ofpockets.
 17. The flow path assembly of claim 16, wherein the outer endof each nozzle airfoil defines an outlet to direct the flow of coolingfluid from the internal cavity to the respective pocket of the pluralityof pockets.
 18. The flow path assembly of claim 15, wherein a pluralityof apertures are defined in the outer wall, each aperture defined froman outer surface of the outer wall to the respective pocket of theplurality of pockets.
 19. The flow path assembly of claim 18, whereinthe flow of cooling fluid is directed through each aperture to theplurality of pockets.
 20. The flow path assembly of claim 15, whereinthe fluid curtain is formed along an outer surface of each nozzleairfoil.